Blade with abrasive tip

ABSTRACT

A blade includes an airfoil section extending between leading and trailing edges, first and second opposed sides each joining the leading and trailing edges, and an inner end and a free end. The blade also includes an abrasive tip at the free end of the airfoil section. The abrasive tip includes particles diposed in a matrix material. The matrix material is a polymeric material that has a glass transition temperature greater than or equal to about 225 degrees C. (487 degrees F.). A gas turbine engine and a method of fabricating a blade are also disclosed.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a continuation of U.S. patent application Ser. No.16/881,358, filed May 22, 2020; the disclosure of which is incorporatedby reference in its entirety herein.

BACKGROUND

This disclosure relates to abrasive tips for rotatable blades. Abradableseals or coatings (rub coatings) can be used to protect moving partsfrom damage during rub interaction while providing a small clearance.Such seals are used in turbomachines to interface with abrasive tips ofa rotating blade stage.

SUMMARY

A blade according to an exemplary embodiment of this disclosure, amongother possible things includes an airfoil section extending betweenleading and trailing edges, first and second opposed sides each joiningthe leading and trailing edges, and an inner end and a free end. Theblade also includes an abrasive tip at the free end of the airfoilsection. The abrasive tip includes particles disposed in a matrixmaterial. The matrix material is a polymeric material that has a glasstransition temperature greater than or equal to about 225 degrees C.(487 degrees F.).

In a further example of the foregoing, the airfoil section comprises analuminum or aluminum-based material.

In a further example of any of the foregoing, the particles include atleast one of alumina (Al2O3), zirconia (ZrO2), oxides, nitrides,carbides, oxycarbides, oxynitrides, diamond and combinations thereof.

In a further example of any of the foregoing, the matrix materialincludes at least one of polyamide, polyimide, bismaleimide, orcombinations thereof.

In a further example of any of the foregoing, the blade includes fibersdisposed in the matrix.

In a further example of any of the foregoing, the fibers are disposed ina proximal area of the abrasive tip and the particles are disposed in adistal area of the abrasive tip.

In a further example of any of the foregoing, the blade includes anadhesive bonding the abrasive tip to the airfoil section.

In a further example of any of the foregoing, the airfoil sectionincludes an overcoat.

In a further example of any of the foregoing, the airfoil sectionincludes a reinforcement at the leading edge.

In a further example of any of the foregoing, the blade is a fan bladefor a gas turbine engine.

A gas turbine engine according to an exemplary embodiment of thisdisclosure, among other possible things includes a compressor section, acombustor in fluid communication with the compressor section; a turbinesection in fluid communication with the combustor, and a fan rotatablycoupled with the turbine section. The fan includes a plurality ofcircumferentially-spaced rotatable blades. Each of the blades include anairfoil section extending between leading and trailing edges, first andsecond opposed sides each joining the leading and trailing edges, and aninner end and a free tip end and an abrasive tip at the free end of eachairfoil section. The abrasive tip includes particles disposed in amatrix material. The matrix material is a polymeric material that has aglass transition temperature greater than or equal to about 225 degreesC. (487 degrees F.). The gas turbine engine also includes a sealcircumscribing the plurality of circumferentially-spaced rotatableblades, the seal being contactable with, and abradable by, the abrasivetip.

In a further example of the foregoing, the matrix material includes atleast one of polyamide, polyimide, bismaleimide, or combinationsthereof.

In a further example of any of the foregoing, the gas turbine engineincludes fibers disposed in the matrix.

In a further example of any of the foregoing, the fibers are disposed ina proximal area of the abrasive tip and the particles are disposed in adistal area of the abrasive tip.

A method of fabricating a blade according to an exemplary embodiment ofthis disclosure, among other possible things includes fabricating anabrasive tip, the abrasive tip comprising particles disposed in a matrixmaterial; and attaching the abrasive tip to a free end of an airfoilsection of a blade after the fabricating.

In a further example of the foregoing, the fabricating includes placingthe matrix material into a mold, placing a layer of the particles intothe mold over the matrix material, and consolidating the matrix materialand the layer of particles such that the matrix material at leastpartially infiltrates the layer of the particles.

In a further example of any of the foregoing, the attaching includesbonding the abrasive tip to the free end of the airfoil section by anadhesive.

In a further example of any of the foregoing, the consolidating includeselevating the temperature of the matrix material and layer of particlesto a first temperature that is higher than a glass transitiontemperature of the matrix material. The bonding includes curing theadhesive at a second temperature. The second temperature is lower thanthe first temperature.

In a further example of any of the foregoing, the fabricating includesforming a prepeg, the prepeg including fibers disposed in the matrixmaterial; placing the prepeg into a mold; placing a layer of theparticles into the mold over the prepeg; and consolidating the prepegand the layer of particles such that the matrix material at leastpartially infiltrates the layer of particles.

In a further example of any of the foregoing, the airfoil sectionincludes at least one of an overcoat and a reinforcement prior to theattaching.

Although the different examples have the specific components shown inthe illustrations, embodiments of this invention are not limited tothose particular combinations. It is possible to use some of thecomponents or features from one of the examples in combination withfeatures or components from another one of the examples.

These and other features disclosed herein can be best understood fromthe following specification and drawings, the following of which is abrief description.

BRIEF DESCRIPTION OF THE DRAWINGS

The various features and advantages of the present disclosure willbecome apparent to those skilled in the art from the following detaileddescription. The drawings that accompany the detailed description can bebriefly described as follows.

FIG. 1 illustrates an example gas turbine engine.

FIG. 2 illustrates an isolated view of the fan section of the gasturbine engine of FIG. 1.

FIG. 3 illustrates an abrasive tip interfacing with an abradable seal.

FIG. 4 illustrates a cross-section of an abrasive tip.

FIG. 5 illustrates a method of fabricating a blade.

FIGS. 6A-B illustrate the blade during fabrication according to themethod of FIG. 5.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. The fan section 22 drivesair along a bypass flow path B in a bypass duct defined within a housing15 such as a fan case or nacelle, and also drives air along a core flowpath C for compression and communication into the combustor section 26then expansion through the turbine section 28. Although depicted as atwo-spool turbofan gas turbine engine in the disclosed non-limitingembodiment, it should be understood that the concepts described hereinare not limited to use with two-spool turbofans as the teachings may beapplied to other types of turbine engines including three-spoolarchitectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects, a first (or low) pressure compressor 44 and a first (orlow) pressure turbine 46. The inner shaft 40 is connected to the fan 42through a speed change mechanism, which in exemplary gas turbine engine20 is illustrated as a geared architecture 48 to drive a fan 42 at alower speed than the low speed spool 30. The high speed spool 32includes an outer shaft 50 that interconnects a second (or high)pressure compressor 52 and a second (or high) pressure turbine 54. Acombustor 56 is arranged in exemplary gas turbine 20 between the highpressure compressor 52 and the high pressure turbine 54. A mid-turbineframe 57 of the engine static structure 36 may be arranged generallybetween the high pressure turbine 54 and the low pressure turbine 46.The mid-turbine frame 57 further supports bearing systems 38 in theturbine section 28. The inner shaft 40 and the outer shaft 50 areconcentric and rotate via bearing systems 38 about the engine centrallongitudinal axis A which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded through the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of the low pressure compressor, or aftof the combustor section 26 or even aft of turbine section 28, and fan42 may be positioned forward or aft of the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five 5:1. Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1 and less than about 5:1. Itshould be understood, however, that the above parameters are onlyexemplary of one embodiment of a geared architecture engine and that thepresent invention is applicable to other gas turbine engines includingdirect drive turbofans. A significant amount of thrust is provided bythe bypass flow B due to the high bypass ratio. The fan section 22 ofthe engine 20 is designed for a particular flight condition—typicallycruise at about 0.8 Mach and about 35,000 feet (10,668 meters). Theflight condition of 0.8 Mach and 35,000 ft (10,668 meters), with theengine at its best fuel consumption—also known as “bucket cruise ThrustSpecific Fuel Consumption (‘TSFC’)”—is the industry standard parameterof lbm of fuel being burned divided by lbf of thrust the engine producesat that minimum point. “Low fan pressure ratio” is the pressure ratioacross the fan blade alone, without a Fan Exit Guide Vane (“FEGV”)system. The low fan pressure ratio as disclosed herein according to onenon-limiting embodiment is less than about 1.45. “Low corrected fan tipspeed” is the actual fan tip speed in ft/sec divided by an industrystandard temperature correction of [(Tram ° R.)/(518.7° R.)]0.5. The“Low corrected fan tip speed” as disclosed herein according to onenon-limiting embodiment is less than about 1150 ft/second (350.5meters/second).

The example gas turbine engine includes the fan section 22 thatcomprises in one non-limiting embodiment less than about 26 fan 42blades 62. In another non-limiting embodiment, the fan section 22includes less than about 20 fan 42 blades 62. Moreover, in one disclosedembodiment the low pressure turbine 46 includes no more than about 6turbine rotors. In another non-limiting example embodiment the lowpressure turbine 46 includes about 3 turbine rotors. A ratio between thenumber of fan 42 blades 62 and the number of low pressure turbine rotorsis between about 3.3 and about 8.6. The example low pressure turbine 46provides the driving power to rotate the fan section 22 and thereforethe relationship between the number of turbine rotors in the lowpressure turbine 46 and the number of blades 62 in the fan section 22disclose an example gas turbine engine 20 with increased power transferefficiency.

FIG. 2 illustrates an isolated view of the fan section 22 of the engine20. Though the description herein is made in reference to the fansection 22, it should be understood that the features described hereincould be used in other parts of the engine 20 as well. The fan 42includes a rotor 60 that has a plurality of circumferentially-spacedblades 62. Each blade 62 includes an airfoil section 64 that extendsbetween leading and trailing edges 66/68, first and second opposed sides70/72 that each joins the leading and trailing edges 66/68, and an innerend 74 and a free tip end 76.

FIG. 3 shows a cutaway view of a representative portion of the airfoilsection 64 of one of the blades 62 and a portion of the abradable seal80. The airfoil section 64 could be formed of a metallic material, suchas aluminum or an aluminum alloy. In other examples, the airfoil section64 could be formed of a ceramic-base composite, such as a carbon fibercomposite. In some examples, the airfoil section 64 includes aprotective overcoat 62 a , such as a polymeric overcoat. For instance,the polymeric overcoat 62 a could protect the underlying airfoil section64 from erosion due to foreign particulate ingested into the engine 20.The overcoat 62 a can be a polyurethane-based coating, an epoxy-basedcoating, or a silicone rubber-based coating, but is not limited to thesetypes of coatings or materials. The overcoat 62 a can cover the firstand second sides 70/72 of the blades 62 and can span the entire lateralsurface of the blade 62 between the leading and trailing edges 66/68.The overcoat 62 a could be bonded to the airfoil section 64 by anadhesive such as an epoxy or epoxy-based adhesive.

The airfoil section 64 can include reinforcements 65 along the leadingedge 66 of the airfoil section. The reinforcements 65 could be made ofmetallic material, such as titanium or titanium-based alloys. Thereinforcements 65 could protect the airfoil section 64 from damage uponencountering a foreign object, for example. Though the examplereinforcements 65 in FIG. 2 are shown along the leading edge 66 of theairfoil section 64, other reinforcements are contemplated. Furthermore,though the example reinforcements 65 in FIG. 2 track along the entireleading edge 66, in other examples, the reinforcements could track alongless than the entire leading edge 66. The reinforcements 65 could bebonded to the airfoil section 64 by an adhesive such as an epoxy orepoxy-based adhesive.

Each blade includes an abrasive tip 78 at the free tip end 76. The fancase 15 is annular in shape and circumscribes the blades 62. The fansection 22 is designed such that the abrasive tips 78 of the blades 62rub against the fan case 15 during rotation. In this regard, the fancase 15 includes an abradable seal 80 mounted on a radially inner sideof the fan case 15. The abradable seal 80 can be formed of apolymeric-based material, such as a polymer matrix composite, in someexamples.

When two components are in rubbing contact, at least one of thecomponents may wear. The term “abradable” refers to the one of the twocomponents that wears, while the other component is “abrasive” and doesnot wear or wears less. Thus, when the abrasive tips 78 of the blades 62rub against the seal 80, the seal 80 will be worn whereas the abrasivetips 78 will not wear or will wear less than the seal 80. The word“abrasive” thus also implies that there is or can be contact with anabradable component.

Friction between a blade tip and a surrounding case generates heat. Theheat can be conducted into the case, into the blade, or both. However,in particular for metal blades 62 and polymeric-based seals 80, themetal of the blade 62 is generally a better thermal conductor than thepolymer of the seal 80, and a majority of the heat thus can conduct intothe blade. While this may normally not present any detriments for aplain metal blade, the heat conduction can be detrimental to a metalblade that has an overcoat 62 a and/or reinforcements 65. The heat cancause softening and/or flow of the adhesives discussed above, which cancause delamination of the polymeric overcoat and thus compromise theerosion protection and/or loosen the bond between the blade 62 andreinforcements 65. Furthermore, some abrasive tips require extremelyhigh processing temperatures during the process of manufacturing theabrasive tip. In this regard, the subsequent disclosure provides anabrasive tip 78 with suitable abrasive properties and improvedtemperature resistance and an improved method of manufacturing theabrasive tip 78.

FIG. 4 illustrates a cross-section of representative portion of anexample abrasive tip 78. The example abrasive tip 78 includes hard“grit” particles 82 in a matrix 84. The abrasive tip 78 is attached tothe airfoil section 64 by an adhesive 85, such as an epoxy orepoxy-based adhesive. Any suitable grit particles 82 known in the artcould be used. For example, the particles 82 could be ceramic-basedparticles such as alumina (Al₂O₃), zirconia (ZrO₂), other oxides,nitrides, carbides, oxycarbides, oxynitrides, diamond and combinationsthereof. The matrix 84 is a polymeric-based material with a high glasstransition temperature, for example, a glass transition temperature ofgreater than or equal to about 225 degrees C. (437 degrees F.). In afurther example, the glass transition temperature of the matrix 84material is greater than or equal to about 275 degrees C. (527 degreesF.). Example matrix 84 materials include polyamides, polyimides,bismaleimide, and combinations thereof. The high glass transitiontemperature of the matrix 84 provides improved heat resistance to theabrasive tip 78. First, the high glass transition temperature improvesretention of the hard particles 82 in the abrasive tip 78 andmaintenance of the thickness of the abrasive tip 78 during operation ofthe engine 20 because the matrix 84 has improved resistance to flow evenunder rubbing conditions, which generate heat as discussed above. Thisin turn contributes to improved longevity of the abrasive tip 78.

Also, the matrix 84 material has low heat conductivity as compared tometal matrix materials. In one example, the heat conductivity of thematrix 84 material is about 2 W/(mK) for temperatures in the range ofabout 150-200 degrees C. (302-392 degrees F.). Metal matrix materialshave heat conductivities for a similar temperature range of about 75W/(mK). Therefore, the heat conductivity of the matrix 84 material isabout an order of magnitude or more lower than metallic matrixabrasives. The low heat conductivity of the matrix 84 material reducesheat transfer to the adhesive 85 and other parts of the blade 62 such asthe airfoil section 64 and the reinforcements 65. This in turn mitigatespossible softening/weakening of the adhesive 85 that bonds the abrasivetip 78 to the airfoil section 64. Other adhesives in other parts of theblade 62, such as adhesives used to bond reinforcements 65 to theairfoil section 64 or adhesives used to bond overcoat 62 a to the blade62, as discussed above, benefit as well.

The hard particles 82 can have an average maximum dimension in aparticle size range of 10-200 micrometers. The hard particles 82 mayprotrude from the matrix 84 or be completely covered by the metalmatrix. In the illustrated example in FIG. 4, the hard particles 82 arefaceted and thus have angled facets 82 a. The angled facets 82 a providerelatively sharp corners that facilitate efficient “cutting” through theabradable seal 80 with low cutting forces, which lowers frictions and,in turn, contributes to lowering the amount of heat generated.

In some examples, the particles 82 are generally situated in a distal(e.g., furthest from the airfoil section 64) area of the abrasive tip78, and the abrasive tip 78 further includes fibers 86 in the matrix 84near a proximal end (e.g. nearest the airfoil section 64) of theabrasive tip 78. The fibers 86 could be carbon, glass, ceramic orpolymeric-based fibers.

The fibers 86 in the example of FIGS. 4 and 6-B (discussed below) arearranged in a woven configuration, but it should be understood thatother configurations, such as unidirectional or random orientations arealso contemplated. The fibers 86 may reinforce the abrasive tip 78during manufacturing, which is discussed in more detail below. Ingeneral, the fibers 86 can provide improved strength for handling of theabrasive tip 78 during fabrication of the abrasive tip 78 and subsequentbonding of the abrasive tip 78 to the airfoil section 64. The fibers 86can be selected to impart certain properties to the abrasive tip 78. Ina particular example, the fibers 86 are carbon-based fibers, which havehigh thermal conductivity in their longitudinal direction which improvesheat dissipation from individual particles 82 that are in rubbingcontact with a mating surface (e.g., seal 80). This can assist inimproving particle 82 retention in the matrix 84. In another example,the fibers 86 are silica-based glass fibers, which have lower thermalconductivity than carbon-based fibers. The lower thermal conductivityreduces heat dissipation to the adhesive 85 and can therefor assist inprotecting the adhesive 85 from excessive heating.

FIG. 5 schematically illustrates a method 500 of fabricating theabrasive tip 78. FIGS. 6A-B schematically illustrate the abrasive tip 78during fabrication according to the method 500. In optional step 502, aprepeg 100 including fibers 86 and matrix 84 material is fabricated. Theprepeg 100 can be fabricated according to any known method, butgenerally includes laying up the fibers 86 into an orientation andinfiltrating the fibers with the matrix material 84.

In step 504, the prepeg 100 is placed in a mold 102 as shown in FIG. 6A.If a prepeg 100 is not being used, matrix material 84 is placed in themold 102 in step 504. A layer 104 of hard particles 82 is placed in themold over the prepeg 100/matrix material 84. The mold 102 can have ageometry that is similar to the final desired geometry of the abrasivetip 78. In other examples, the mold 102 can be used to form a largerproduct which can be machined to shape after the method 500 as would beknown in the art.

In step 506, the material in the mold 102 is consolidated to form anabrasive tip 78 as shown in FIG. 6B. The consolidation can include theapplication of heat and/or pressure, in one example. For instance, theconsolidation can include heating the matrix material 84 to atemperature above it glass transition temperature. In another example,the consolidation can include inducing vacuum in the mold. In general,the consolidation causes the matrix material 84 to partially or fullyinfiltrate the layer 104 of hard particles 82. The consolidation alsocauses consolidation of the prepeg 100, if being used.

In step 508, the abrasive tip 78 is removed from the mold. As discussedabove, depending on the shape of the mold 102, the abrasive tip 78 maybe machined to a desired shape after step 508.

In step 510, the abrasive tip 78 is attached to the airfoil section 64.For example, the abrasive tip 78 is bonded to the airfoil section 64 byadhesive 85, which could include applying the adhesive 85 to theabrasive tip 78 and/or airfoil section 64 and curing the adhesive 85 byany known method. The bonding can include priming steps that improve thebond of the adhesive 85, such as application of a primer material or anyother priming steps as would be known in the art. The bonding includescuring the adhesive 85. In some examples, the airfoil section 64 alreadyincludes the overcoat 62 a and/or the reinforcements 65 during step 510.In this regard, the overcoat 62 a and/or reinforcements 65 are notsubjected to the consolidation step 506, which in some examples couldrequire elevated temperatures. Though heating could be used to cure theadhesive 85, in general, curing the adhesive 85 can be effectuated atlower temperatures than the consolidating step 506 discussed above.Therefore, the adhesives that bond the overcoat 62 a/reinforcements 65to the airfoil section 64 are not subjected to elevated temperatureswhich could cause softening or delamination as discussed above.

As discussed above, optional fibers 86 can reinforce the abrasive tip 78to provide improved strength for handling of the abrasive tip 78 duringsteps 508 and 510.

Though the foregoing method 500 was described with respect tomanufacture of the abrasive tip 78,

Although a combination of features is shown in the illustrated examples,not all of them need to be combined to realize the benefits of variousembodiments of this disclosure. In other words, a system designedaccording to an embodiment of this disclosure will not necessarilyinclude all of the features shown in any one of the Figures or all ofthe portions schematically shown in the Figures. Moreover, selectedfeatures of one example embodiment may be combined with selectedfeatures of other example embodiments.

The preceding description is exemplary rather than limiting in nature.Variations and modifications to the disclosed examples may becomeapparent to those skilled in the art that do not necessarily depart fromthe essence of this disclosure. The scope of legal protection given tothis disclosure can only be determined by studying the following claims.

What is claimed is:
 1. A blade comprising: an airfoil section extendingbetween leading and trailing edges, first and second opposed sides eachjoining the leading and trailing edges, and an inner end and a free end;and an abrasive tip at the free end of the airfoil section, wherein theabrasive tip includes particles disposed in a matrix material, andwherein the matrix material is a polymeric material that has a glasstransition temperature greater than or equal to about 225 degrees C.(487 degrees F.).
 2. The blade as recited in claim 1, wherein theairfoil section comprises an aluminum or aluminum-based material.
 3. Theblade as recited in claim 1, wherein the particles include at least oneof alumina (Al₂O₃), zirconia (ZrO₂), oxides, nitrides, carbides,oxycarbides, oxynitrides, diamond and combinations thereof.
 4. The bladeas recited in claim 1, wherein the matrix material includes at least oneof polyamide, polyimide, bismaleimide, or combinations thereof.
 5. Theblade as recited in claim 1, further comprising fibers disposed in thematrix.
 6. The blade as recited in claim 5, wherein the fibers aredisposed in a proximal area of the abrasive tip and the particles aredisposed in a distal area of the abrasive tip.
 7. The blade as recitedin claim 1, further comprising an adhesive bonding the abrasive tip tothe airfoil section.
 8. The blade as recited in claim 1, wherein theairfoil section includes an overcoat.
 9. The blade as recited in claim1, wherein the airfoil section includes a reinforcement at the leadingedge.
 10. The blade as recited in claim 1, wherein the blade is a fanblade for a gas turbine engine.
 11. A gas turbine engine comprising: acompressor section; a combustor in fluid communication with thecompressor section; a turbine section in fluid communication with thecombustor; a fan rotatably coupled with the turbine section, the fanincluding a plurality of circumferentially-spaced rotatable blades, eachof the blades including: an airfoil section extending between leadingand trailing edges, first and second opposed sides each joining theleading and trailing edges, and an inner end and a free tip end, and anabrasive tip at the free end of each airfoil section, wherein theabrasive tip includes particles disposed in a matrix material, andwherein the matrix material is a polymeric material that has a glasstransition temperature greater than or equal to about 225 degrees C.(487 degrees F.); and a seal circumscribing the plurality ofcircumferentially-spaced rotatable blades, the seal being contactablewith, and abradable by, the abrasive tip.
 12. The gas turbine engine asrecited in claim 11, wherein the matrix material includes at least oneof polyamide, polyimide, bismaleimide, or combinations thereof.
 13. Thegas turbine engine as recited in claim 11, further comprising fibersdisposed in the matrix.
 14. The gas turbine engine as recited in claim13, wherein the fibers are disposed in a proximal area of the abrasivetip and the particles are disposed in a distal area of the abrasive tip.15. A method of fabricating a blade, comprising: fabricating an abrasivetip, the abrasive tip comprising particles disposed in a matrixmaterial; attaching the abrasive tip to a free end of an airfoil sectionof a blade after the fabricating.
 16. The method of claim 15, whereinthe fabricating includes: placing the matrix material into a mold;placing a layer of the particles into the mold over the matrix material;and consolidating the matrix material and the layer of particles suchthat the matrix material at least partially infiltrates the layer of theparticles.
 17. The method of claim 16, wherein the attaching includesbonding the abrasive tip to the free end of the airfoil section by anadhesive.
 18. The method of claim 17, wherein the consolidating includeselevating the temperature of the matrix material and layer of particlesto a first temperature that is higher than a glass transitiontemperature of the matrix material, and wherein the bonding includescuring the adhesive at a second temperature, wherein the secondtemperature is lower than the first temperature.
 19. The method of claim15, wherein the fabricating includes: forming a prepeg, the prepegincluding fibers disposed in the matrix material; placing the prepeginto a mold; placing a layer of the particles into the mold over theprepeg; and consolidating the prepeg and the layer of particles suchthat the matrix material at least partially infiltrates the layer ofparticles.
 20. The method of claim 15, wherein the airfoil sectionincludes at least one of an overcoat and a reinforcement prior to theattaching.